摘要
以全流量补燃循环氢氧发动机系统为研究对象,对其的动态响应特性进行了研究。建立了描述全流量补燃循环发动机动态特性的非线性数学模型,将免疫策略算法同龙格-库塔法结合起来,提出了求解含有隐式项的常微分方程组的变步长龙格-库塔方法,并应用该方法对全流量补燃循环发动机系统的动态响应特性进行了仿真计算。计算结果表明,当发动机在某一个稳定工况工作时,发动机入口推进剂压强的变化对发动机性能参数的影响不大,发动机的参数都能比较平稳的过渡到一个新的稳定工况;当发动机在短时间内进行小范围的推力调节时,发动机参数的过渡过程的曲线也比较平稳,但是当在短时间内进行大范围推力调节时,参数的过渡过程的曲线振动比较剧烈,因此应当在进行大范围推力调节时,应当对调节时间进行适当延长,或者分级进行推力调节。
Using the full-flow staged combustion cycle LH2/LOX engine as an example,the engine response characteristics were investigated. The non-linear dynamic mathematical model of the full-flow staged combustion cycle was presented. The variable step Runge-Kutta algorithm for solving the ordinary differential equations was developed. The immunity strategy algorithm and Runge-Kutta algorithm were combined to form this new algorithm. Applying the variable step Runge-Kutta algorithm,the transient performance of the full-flow staged combustion cycle was simulated. Calculating results show that variation of the propellant inlet pressure makes little impact on engine performance parameters,and the engine will transit to a new stable state. The parameter transition curves are smooth. When the engine thrust is adjusted in a small range within a short period,the parameter transition curves are also reposeful. But when the adjusting range becomes large, the parameter transition curves will fluctuate acutely. Thus ,when large range adjusting is required ,the time of adjusting should be lengthened or the adjusting should be divided into adequate steps.
出处
《航空动力学报》
EI
CAS
CSCD
北大核心
2006年第2期410-416,共7页
Journal of Aerospace Power
关键词
航空
航天推进系统
液体推进剂火箭发动机
瞬变特性
常微分方程组
干扰因素
全流量补燃
aerospace propulsion system
liquid propellant rocket engine
transient performance
ordinary differential equations
interfering factor
full-flow staged combustion cycle
